1. Field of the Invention
This invention relates to gas turbine engines and more particularly to the cooling of turbine components in engines having high turbine inlet temperatures.
2. Description of the Prior Art
A gas turbine engine principally includes a compressor section, a combustion section and a turbine section. Each of the sections is individually designed, where possible, to maximize local efficiencies and improve the associated overall performance of the engine. Care must be exercised, however, to ensure that local efficiency improvements do not disturb the operating compatibility between the various engine components of the principal sections.
The combustion section is positioned immediately downstream of the compressor section and receives high pressure gases which comprise the engine working medium discharged from the compressor. A portion of the working medium gases is mixed with fuel in a combustion chamber to form a combustible mixture which is burned to increase the kinetic energy of the flowing gases. To decrease the amount of unburned hydrocarbons which are discharged from the combustion chamber and to improve chamber performance, it is desired to burn the combustible mixture at high temperatures. Additional combustion chamber performance increases are achieved by minimizing the pressure loss of the working medium as the medium flows through the combustion section. It is, therefore, well known in the art that a combustion chamber having maximum temperatures and minimum flow losses offers the highest local benefits.
From the combustion section the working medium is flowed to the turbine where an annular nozzle comprising a plurality of turbine vanes directs the working medium gases, which include the hot combustion gases produced in the upstream chamber, at a preferred angle into a row of downstream turbine blades. The blades receive the working medium and extract kinetic energy from the medium to drive the compressor and engine accessories. Although the blades and vanes are fabricated from the finest high temperature materials in order to survive the hostile environment of the combustion gases, the local temperatures of the combustion gases still exceed the maximum acceptable temperatures for all known suitable blade and vane materials.
It is widely known that blade and vane temperatures can be held within safe limits by flowing cooling air over the internal and external surfaces of the airfoil sections. Compressor exit air is the most highly pressurized air available within the engine and is utilized for cooling. In one typical embodiment cooling air is flowed from a port at the inner diameter of the compressor gas path through various conduit means to the turbine section of the engine. The pressure of the cooling air which flows to the turbine is reduced by flow losses through the conduit means and is typically at a value of 95% to 96% of the compressor exit pressure by the time the cooling air reaches the airfoil section. As long as the pressure loss of the working medium flowing through the combustion section is higher than the pressure loss of the cooling air through the various conduit means, a positive flow of cooling air through the blades and vanes is maintained.
As has been discussed above, minimization of the combustion section pressure loss is one of the design goals toward achieving optimum overall engine performance. Although current commercial engines exhaust combustion gases to the turbine at approximately 94% of the compressor exit pressure, modern combustion chambers exhausting gases at approximately 96% of the compressor exit pressure have been designed. The minimized pressure loss combustor, while remaining highly desirable, is somewhat limited in utility by the downstream components of the turbine which require a positive flow of cooling air. U.S. Pat. No. 3,628,880 to Smuland shows a turbine vane having intricate cooling systems to effect acceptable vane cooling in the destructive turbine environment of an engine having high turbine inlet temperatures. Various combinations of impingement, convective and film cooling maintain the metal temperature at or below acceptable levels as long as an adequate supply of cooling air is provided. Impingement cooling which is provided at the interior walls of the Smuland airfoil, requires a high pressure differential between the cooling air and the turbine working medium and is especially sensitive to reduced cooling air pressure. In normal operation cooling air is accelerated through small diameter holes in a baffle insert to a velocity at which air impinges on the interior walls of the airfoil. If the pressure drop across the baffle is not high enough low velocity flow occurs and impingement cooling is not accomplished. In a typical modern engine the pressure drop across the leading edge of the foil is within a range of three to six psi at take-off, and somewhat lower at altitude conditions. Inasmuch as pressure drops substantially greater than that are required for impingement cooling, the cooling flow is exhausted to a lower pressure along the suction side of the airfoil section rather than to the leading edge. It is known in the art that film cooling of the leading edge offers an attractive and efficient means for cooling the leading edge. In current engines the capacity for film cooling in conjunction with impingement cooling at the leading edge is severely limited by the inability to exhaust impingement cooling flow at the leading edge of the airfoil as described above.
Continuing efforts are underway to maximize the pressure of the cooling air in the vane region in order to ensure an adequate pressure differential between the cooling air and the working medium gases to which the cooling air is exhausted. Improved turbine cooling is of increased importance as efficiently operating low pressure drop combustion chambers become available.